Molybdenum-silicon-boron with noble metal barrier layer

ABSTRACT

An article includes a substrate formed of a molybdenum-based alloy. A barrier layer is disposed on the substrate. The barrier layer is formed of at least one noble metal.

CROSS-REFERENCED TO RELATED APPLICATION

This application is a continuation of U.S. patent application Ser. No.16/432,417 filed Jun. 5, 2019, which is a continuation of U.S. patentapplication Ser. No. 15/149,344 filed May 9, 2016, now U.S. Pat. No.10,329,926 issued Jun. 25, 2019, and both of which are hereinincorporated by reference in their entireties.

BACKGROUND

Gas turbine engines and other machines often operate under severelyelevated temperature conditions. Components that are exposed to theelevated temperatures can be formed of temperature-resistant metalalloys, such as nickel alloys or molybdenum alloys. Molybdenum alloysgenerally have good strength and other properties at high temperaturesbut are potentially susceptible to oxidation in high temperatureoxidizing environments. The inclusion of silicon and boron in the alloycan reduce oxygen infiltration into the alloy through formation of apassive borosilicate scale.

SUMMARY

An article according to an example of the present disclosure includes asubstrate formed of a molybdenum-based alloy. The molybdenum-based alloyhas a composition that has molybdenum, silicon, and boron. A barrierlayer is disposed on the substrate. The barrier layer is formed of atleast one noble metal.

A further embodiment of any of the foregoing embodiments includes ametal interlayer disposed between the substrate and the barrier layer.

In a further embodiment of any of the foregoing embodiments, the metalinterlayer is formed of a different metal than the barrier layer and isselected from the group consisting of copper, gold, platinum, andcombinations thereof.

In a further embodiment of any of the foregoing embodiments, the metalinterlayer includes copper.

In a further embodiment of any of the foregoing embodiments, the atleast one noble metal is selected from the group consisting ofruthenium, rhodium, palladium, silver, osmium, iridium, platinum, gold,and combinations thereof.

In a further embodiment of any of the foregoing embodiments, at leastone noble metal includes platinum.

A further embodiment of any of the foregoing embodiments includes atopcoat disposed on the barrier layer.

In a further embodiment of any of the foregoing embodiments, the topcoatis formed of a silica material.

In a further embodiment of any of the foregoing embodiments, thesubstrate defines an internal cavity, and the barrier layer is disposedon the substrate in the internal cavity.

A method of fabricating an article according to an example of thepresent disclosure includes depositing a barrier layer onto a substrateformed of a molybdenum-based alloy. The molybdenum-based alloy has acomposition that includes molybdenum, silicon, and boron. The barrierlayer has at least one noble metal.

A further embodiment of any of the foregoing embodiments includes, priorto depositing the barrier layer, depositing a metal interlayer onto thesubstrate.

In a further embodiment of any of the foregoing embodiments, the metalinterlayer is formed of a different metal than the barrier layer and isselected from the group consisting of copper, gold, platinum, andcombinations thereof.

A further embodiment of any of the foregoing embodiments includesthermally treating the substrate with the metal interlayer and thebarrier layer, and conducting the thermal treating at a temperatureabove the melting point of the metal of the metal interlayer but belowthe melting point of the at least one noble metal.

A further embodiment of any of the foregoing embodiments includesdepositing the metal interlayer and depositing the barrier layer usingelectroplating.

In a further embodiment of any of the foregoing embodiments, thesubstrate defines an internal cavity, and the barrier layer is depositedon the substrate in the internal cavity.

A gas turbine engine according to an example of the present disclosureincludes an engine section selected from the group consisting of acompressor section, a combustor section, and a turbine section. Theengine section has a turbine engine component including a substrateformed of a molybdenum-based alloy. The molybdenum-based alloy has acomposition that includes molybdenum, silicon, and boron. A barrierlayer is disposed on the substrate. The barrier layer is formed of atleast one noble metal.

A further embodiment of any of the foregoing embodiments includes ametal interlayer disposed between the substrate and the barrier layer.

In a further embodiment of any of the foregoing embodiments, metalinterlayer is formed of a different metal than the barrier layer and isselected from the group consisting of copper, gold, platinum, andcombinations thereof.

In a further embodiment of any of the foregoing embodiments, at leastone noble metal is selected from the group consisting of ruthenium,rhodium, palladium, silver, osmium, iridium, platinum, gold, andcombinations thereof.

In a further embodiment of any of the foregoing embodiments, substratedefines an internal cavity, and the metal interlayer and the barrierlayer are disposed on the substrate in the internal cavity.

BRIEF DESCRIPTION OF THE DRAWINGS

The various features and advantages of the present disclosure willbecome apparent to those skilled in the art from the following detaileddescription. The drawings that accompany the detailed description can bebriefly described as follows.

FIG. 1 illustrates an example gas turbine engine.

FIG. 2 illustrates a cross-section of a portion of an article that has aMo—Si—B alloy substrate and a noble metal barrier layer.

FIG. 3 illustrates a cross-section of a portion of an article that has aMo—Si—B alloy substrate, a noble metal barrier layer, and a metalinterlayer.

FIG. 4 illustrates a cross-section of a portion of an article that has aMo—Si—B alloy substrate, a noble metal barrier layer, a metalinterlayer, and a topcoat.

FIG. 5 illustrates an article with a substrate that has an internalcavity.

FIG. 6 illustrates a cross-section of a portion of the article of FIG.5, with a noble metal barrier layer disposed in the internal cavity.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative enginedesigns can include an augmentor section (not shown) among other systemsor features.

The fan section 22 drives air along a bypass flow path B in a bypassduct defined within a nacelle 15, while the compressor section 24 drivesair along a core flow path C for compression and communication into thecombustor section 26 then expansion through the turbine section 28.Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, the examples herein are not limitedto use with two-spool turbofans and may be applied to other types ofturbomachinery, including direct drive engine architectures, three-spoolengine architectures, and ground-based turbines.

The engine 20 generally includes a low speed spool 30 and a high speedspool 32 mounted for rotation about an engine central longitudinal axisA relative to an engine static structure 36 via several bearing systems38. It should be understood that various bearing systems 38 at variouslocations may alternatively or additionally be provided, and thelocation of bearing systems 38 may be varied as appropriate to theapplication.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a first (or low) pressure compressor 44 and afirst (or low) pressure turbine 46. The inner shaft 40 is connected tothe fan 42 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48, to drivethe fan 42 at a lower speed than the low speed spool 30.

The high speed spool 32 includes an outer shaft 50 that interconnects asecond (or high) pressure compressor 52 and a second (or high) pressureturbine 54. A combustor 56 is arranged between the high pressurecompressor 52 and the high pressure turbine 54. A mid-turbine frame 57of the engine static structure 36 is arranged generally between the highpressure turbine 54 and the low pressure turbine 46. The mid-turbineframe 57 further supports the bearing systems 38 in the turbine section28. The inner shaft 40 and the outer shaft 50 are concentric and rotatevia bearing systems 38 about the engine central longitudinal axis A,which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present invention isapplicable to other gas turbine engines, including direct driveturbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, withthe engine at its best fuel consumption—also known as “bucket cruiseThrust Specific Fuel Consumption (‘TSFC’)”—is the industry standardparameter of 1 bm of fuel being burned divided by 1 bf of thrust theengine produces at that minimum point. “Low fan pressure ratio” is thepressure ratio across the fan blade alone, without a Fan Exit Guide Vane(“FEGV”) system. The low fan pressure ratio as disclosed hereinaccording to one non-limiting embodiment is less than about 1.45. “Lowcorrected fan tip speed” is the actual fan tip speed in ft/sec dividedby an industry standard temperature correction of [(Tram ° R)/(518.7 °R)]^(0.5). The “Low corrected fan tip speed” as disclosed hereinaccording to one non-limiting embodiment is less than about 1150ft/second.

FIG. 2 illustrates a representative portion of an example article 60.For instance, the article 60 is, or is a portion of, a gas turbineengine article in the gas turbine engine 20. For example, the article 60may be in the compressor section 24, the combustor section 26, and/orthe turbine section 28. In this regard, the article 60 may be, but isnot limited to, a blade outer air seal (represented at 60 a and 60 b inFIG. 1), an airfoil (e.g., a vane, represented at 60 c in FIG. 1), ahousing or casing (represented at 60 d in FIG. 1), a structuralcomponent, or the like that is exposed to relatively hot temperatures inthe engine 20 (e.g., in the core gas path C).

The article 60 includes a substrate 62 that is formed of amolybdenum-based alloy. For example, the substrate 62 is the structuralbody of the article 60 that provides the shape of the article 60. Themolybdenum-based alloy has a composition that includes molybdenum,silicon, and boron. For example, the composition of the molybdenum-basedalloy is predominantly molybdenum, and the silicon, boron, and any otheralloy elements are present in substantially lesser amounts than themolybdenum. Alloying elements are generally present in amounts no lessthan 0.2% by weight.

While the silicon and the boron in the alloy can reduce oxygeninfiltration into the alloy through formation of a passive borosilicatescale, in some oxidizing environments the oxygen can permeate the scaleor infiltrate the alloy prior to scale formation and volatilize themolybdenum. The volatilization of molybdenum potentially results inalloy weight loss known as “recession,” which potentially leads to lossof mechanical properties of the alloy. In this regard, the article 60includes a coating 64 that serves as an oxygen barrier to enhanceresistance to oxygen infiltration, oxidation, and volatilization ofmolybdenum.

In the illustrated example, the coating 64 includes a barrier layer 66that is disposed on the substrate 62. For instance, the barrier layer 66is immediately adjacent to, and contiguous with, the substrate 62.Alternatively, additional layers could be provided between the substrate62 and the barrier layer 66 and/or on top of the barrier layer 66.

The barrier layer 66 is formed of at least one noble metal that resistsoxidation and limits oxygen infiltration into the substrate 64. As anexample, the barrier layer 66 has a thickness in the range ofapproximately 1 micrometer to approximately 50 micrometers. Thicknessestoward the lower end of the range may have lower barrier properties buthigher spall-resistance to thermal stress; while thicknesses toward thehigher end of the range may have higher barrier properties but lowerspall-resistance to thermal stresses.

The noble metal or metals of the barrier layer 66 are in metallic form.The term “metallic form” refers to a metallically bonded metal, ratherthan a metal that is ionically or covalently bonded to non-metal atomsin compounds. For example, the noble metal or metals are selected fromruthenium, rhodium, palladium, silver, osmium, iridium, platinum, gold,and combinations thereof. In a further example, the barrier layer 66 isformed essentially or only of a single type of noble metal, and thenoble metal is platinum.

FIG. 3 illustrates another example article 160. In this disclosure, likereference numerals designate like elements where appropriate andreference numerals with the addition of one-hundred or multiples thereofdesignate modified elements that are understood to incorporate the samefeatures and benefits of the corresponding elements. In this example,the article 160 has a coating 164 which, in addition to the barrierlayer 66, includes a metal interlayer 168 that is disposed between thesubstrate 62 and the barrier layer 66. The metal interlayer 168 isformed of a different metal than the barrier layer 66 and is selectedfrom copper, gold, platinum, and combinations thereof. In a furtherexample, the metal interlayer 168 is formed essentially or only of asingle type of metal, and the metal is copper.

The metal interlayer 168 may serve to facilitate bonding of the barrierlayer 66 to the substrate 62. Additionally or alternatively, the metalinterlayer 168 may serve as fluxing agent to promote the formation ofthe borosilicate scale in the molybdenum-based alloy.

FIG. 4 illustrates another example article 260. In this example, thearticle 260 has a coating 264 which, in addition to the barrier layer 66and the metal interlayer 168, includes a topcoat 270 that is disposed onthe barrier layer 66. For example, the topcoat 270 may be formed of, butis not limited to, a silica material. In a further example, the topcoatincludes the silica material and also molybdenum silicide (Mo₃Si).

The topcoat 270 may serve as an additional oxygen barrier layer and/orto protect the barrier layer 66. The topcoat 270 and barrier layer 66may also function cooperatively for stress management. For instance, thetopcoat 270 may be brittle in comparison to the barrier layer 66. Butfor the barrier layer 66, any cracking in the topcoat 270 couldpropagate into the substrate 62. However, the barrier layer 66 isductile in comparison to the topcoat 270 and can potentially arrestcracks that propagate from the topcoat 270. Thus, the barrier layer 66may also serve to enhance durability.

FIG. 5 illustrates a portion of another example article 360. Theillustrated portion is an airfoil but alternatively may be a blade outerair seal or other type of component. The article 360 includes ininternal cavity 372. In this example, the internal cavity 360 is aninternal cooling passage that opens to the exterior surface of thearticle. FIG. 6 illustrates a representative portion of the internalcavity 372. In this example, the substrate 62 defines the internalcavity 372, and the barrier layer 66, metal interlayer 168, and topcoat270 are disposed on the substrate 62 in the internal cavity 372.

The article 60/160/260/360 can be fabricated by depositing the barrierlayer 66 onto the substrate 62. For instance, the noble metal isdeposited onto the substrate 62 using an electroplating process. Ifused, the metal interlayer 168 is deposited onto the substrate 62 priorto deposition of the barrier layer 66. The metal interlayer 168 can alsobe deposited using an electroplating process. Other depositiontechniques may alternatively be used, such as but not limited to,physical vapor deposition, chemical vapor deposition, and plasma spray.However, such processes are typically line-of-sight. To deposit on theinternal cavity 372 or to completely coat all surfaces of the substrate62, the electroplating process may be used.

If the metal interlayer 168 is used, the substrate 62 with the depositedmetal interlayer 168 and the deposited barrier layer 66 may be thermallytreated. For example, the thermal treatment may be conducted undervacuum at a temperature that is above the melting point of the metal ofthe metal interlayer 168 but below the melting point of the noble metalof the barrier layer 66. The heat treatment melts the metal of the metalinterlayer 168 and “brazes” the barrier layer 66 to the substrate 62.The heat treatment may also serve to densify or consolidate the metalinterlayer 168, the barrier layer 66, or both, which may further enhancebarrier properties to oxygen infiltration.

Although a combination of features is shown in the illustrated examples,not all of them need to be combined to realize the benefits of variousembodiments of this disclosure. In other words, a system designedaccording to an embodiment of this disclosure will not necessarilyinclude all of the features shown in any one of the Figures or all ofthe portions schematically shown in the Figures. Moreover, selectedfeatures of one example embodiment may be combined with selectedfeatures of other example embodiments.

The preceding description is exemplary rather than limiting in nature.Variations and modifications to the disclosed examples may becomeapparent to those skilled in the art that do not necessarily depart fromthis disclosure. The scope of legal protection given to this disclosurecan only be determined by studying the following claims.

What is claimed is:
 1. An article comprising: a substrate formed of amolybdenum-based alloy; and a barrier layer immediately adjacent to andcontiguous with the substrate, the barrier layer formed of at least onenoble metal.
 2. The article as recited in claim 1, wherein the at leastone noble metal is selected from the group consisting of ruthenium,rhodium, palladium, silver, osmium, iridium, platinum, gold, andcombinations thereof.
 3. The article as recited in claim 1, wherein theat least one noble metal includes platinum.
 4. The article as recited inclaim 1, further comprising a topcoat disposed on the barrier layer. 5.The article as recited in claim 4, wherein the topcoat is formed of asilica material.
 6. The article as recited in claim 1, wherein thesubstrate defines an internal cavity, and the barrier layer is disposedon the substrate in the internal cavity.
 7. The article as recited inclaim 1, wherein the barrier layer has a thickness of between about 1micrometer and 50 micrometers.
 8. The article as recited in claim 1,wherein the molybdenum-based alloy includes less than about 0.2% byweight of alloying elements, and the balance molybdenum.
 9. The articleas recited in claim 8, wherein the alloying elements include silicon andboron.
 10. A method of fabricating an article, the method comprising:depositing a barrier layer onto a substrate formed of a molybdenum-basedalloy, the barrier layer having at least one noble metal.
 11. The methodas recited in claim 10, wherein the depositing is by electroplating. 12.The method as recited in claim 10, wherein the substrate defines aninternal cavity, and the barrier layer is deposited on the substrate inthe internal cavity.
 13. The method as recited in claim 10, wherein thebarrier layer has a thickness of between about 1 micrometer and 50micrometers after the depositing.
 14. The method as recited in claim 10,wherein the molybdenum-based alloy includes less than about 0.2% byweight of alloying elements, and the balance molybdenum.
 15. The methodas recited in claim 14, wherein the alloying elements include siliconand boron.
 16. A gas turbine engine comprising: an engine sectionselected from the group consisting of a compressor section, a combustorsection, and a turbine section, the engine section having a turbineengine component including, a substrate formed of a molybdenum-basedalloy, and a barrier layer disposed on the substrate, the barrier layerformed of at least one noble metal.
 17. The gas turbine engine asrecited in claim 16, wherein the at least one noble metal is selectedfrom the group consisting of ruthenium, rhodium, palladium, silver,osmium, iridium, platinum, gold, and combinations thereof.
 18. The gasturbine engine as recited in claim 16, wherein the substrate defines aninternal cavity, and the metal interlayer and the barrier layer aredisposed on the substrate in the internal cavity.
 19. The gas turbineengine as recited in claim 16, wherein the barrier layer has a thicknessof between about 1 micrometer and 50 micrometers after the depositing.20. The method as recited in claim 10, wherein the molybdenum-basedalloy includes less than about 0.2% by weight of alloying elements, andthe balance molybdenum, and wherein the alloying elements includesilicon and boron.